This invention relates generally to gas turbine engines, and specifically to cooling techniques for a turbine exhaust assembly. In particular, the invention concerns impingement cooling of a low-bypass turbofan exhaust case (TEC) assembly, with applications in military aviation and high-performance aircraft.
Standard gas turbine engines are built around a power core comprising compressor, combustor and turbine sections, which are arranged in flow series with an upstream inlet and a downstream exhaust nozzle. The compressor compresses air from the inlet. The compressed air is used an oxidant in the combustor, and, in some applications, for accessory pneumatic functions and environmental control. Fuel is injected into the combustor, where it combines with the compressed air and ignites to produce hot combustion gases. The hot combustion gases drive the engine by expansion in the turbine section, and are exhausted to the nozzle through a turbine exhaust case (TEC) assembly.
The turbine section drives the compressor via a rotating shaft, or, in most larger-scale applications, via a number of coaxially nested shafts and independently rotating turbine/compressor assemblies or spools. Each spool, in turn, employs a number of stages, in which rotating blades coupled to the shaft are alternated with stationary vanes coupled to a shroud or other fixed component of the engine housing.
Energy that is not used to drive the compressor and accessory functions is available for extraction and use. In ground-based applications, energy is typically delivered in the form of rotational motion, which is used to drive an electrical generator or other mechanical load coupled to the shaft. In aviation applications, the gas turbine engine also provides reactive thrust.
The relative contributions of rotation and thrust depend upon engine design. In turbojet engines, for example, which are an older design, essentially all the net thrust is generated in the exhaust. In modern turbofan engines, on the other hand, the shaft is used to drive a ducted propeller or forward fan, which generates additional thrust by forcing air through a bypass flow duct surrounding the engine core.
Turbofan engines include low-bypass turbofans, in which the bypass flow is relatively small with respect to the core flow, and high-bypass turbofans, in which the bypass flow is greater. High-bypass turbofans tend to be quieter, cooler and more energy efficient, particularly in subsonic flight applications for commercial and other general-purpose aircraft. Low-bypass turbofans can be somewhat louder and less fuel efficient, but provide greater specific thrust. For these and other reasons, low-bypass turbofans are generally utilized in military jet fighters and other high-performance supersonic aircraft.
In supersonic applications, the turbofan engine is typically provided with an afterburner. Afterburning systems provide thrust augmentation by injecting additional fuel into an augmentor assembly, downstream of the TEC, where it mixes with the core flow and ignites to increase the thrust. Afterburning substantially enhances engine performance, but is also associated with additional costs in efficiency, noise output and thermal signature.
The main design goals for aviation-based gas turbine engines are performance, efficiency, reliability and service life. Performance and efficiency both favor higher combustion temperatures, which increase the engine's specific thrust and overall thermodynamic efficiency. Unfortunately, higher combustion temperatures also result in increased thermal and mechanical loads, particularly for engine components along the hot gas flowpath, downstream of the combustor. This can affect service life and reliability, and increase operational costs associated with maintenance and part replacement.
In high-performance (low-bypass) turbofans, gas path temperatures are often a factor at the TEC assembly, where hot combustion gases flow from turbine section (upstream of the TEC) toward the afterburner/augmentor (downstream of the TEC). The issue can be problematic proximate the forward outer diameter ring (FODR), on the upstream end of the TEC assembly. In this region, operational conditions can sometimes establish a negative pressure differential between the FODR plenum, which surrounds the FODR, and the exhaust gas flow, inside the FODR.
Negative FODR plenum overpressure allows hot gas inflow, impairing cooling efficiency. The result is decreased service life and increased risk of mechanical failure. There is thus a need for improved TEC assembly cooling techniques that provide increased service life and reliability without sacrificing performance and efficiency.